US 8131408 B2 Abstract The present invention provides a method and a calculation system for an aircraft, with at least one sensor for detecting aeroelastic and flight-mechanical momenta of the aircraft, for detecting positions and movements of control surfaces of the aircraft or for detecting speeds of gusts of wind acting on the aircraft and comprising a calculation unit which calculates characteristic quantities of passenger comfort and cabin safety as well as momenta of the aircraft as a function of the sensor data provided by the sensors and a non-linear simulation model of the aircraft.
Claims(17) 1. A system for control, monitoring, and design validation of an aircraft, comprising:
(a) at least one sensor for detecting at least one of aeroelastic and flight-mechanical momenta of the aircraft, positions and movements of control surfaces of the aircraft, and speeds of gusts of wind acting on the aircraft;
(b) a calculation unit having at least one microprocessor, the calculation unit being configured for calculating characteristic quantities of passenger comfort and cabin safety and momenta of the aircraft from data provided by the at least one sensor using a stored non-linear simulation model; and
(c) a control unit configured for controlling the aircraft and coupled to a comparison unit configured for calculating a difference between precalculated quantities and the quantities calculated by the calculation unit.
2. The system according to
M{umlaut over (x)}+D{dot over (x)}+Kx+Fg(x,{dot over (x)},p,t)=p, where
M is a mass matrix,
D is a damping matrix,
K is a rigidity matrix,
x is a hyper-movement vector of the aircraft,
g is the non-linearity vector,
p is a hyper-input vector of the aircraft, and
F is the effectiveness matrix,
wherein the effectiveness matrix describes non-linear characteristics of characteristic quantities which have flight-mechanical characteristic quantities, characteristic quantities of an on-board system and characteristic quantities of the aeroelastics.
3. The system according to
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15. A method for controlling an aircraft, comprising the following steps:
(a) detecting aeroelastic and flight-mechanical momenta of the aircraft, of positions and movements of control surfaces of the aircraft, and speeds of gusts of wind acting on the aircraft, to generate sensor data;
(b) calculating characteristic quantities of passenger comfort and momenta of the aircraft as a function of the generated sensor data and a stored non-linear simulation model of the aircraft; and
(c) calculating a difference between precalculated quantities and the quantities calculated by the calculation unit, the aircraft being controlled as a function of the calculated difference.
16. A computer program product comprising program commands that when executed by a processor of a computer control the computer to perform steps comprising:
detecting aeroelastic and flight-mechanical momenta of an aircraft, positions and movements of control surfaces of the aircraft, and speeds of gusts of wind acting on the aircraft, to generate sensor data;
calculating characteristic quantities of passenger comfort and momenta of the aircraft as a function of the generated sensor data and a stored non-linear simulation model of the aircraft, and
calculating a difference between precalculated quantities and the quantities calculated by the calculation unit, the aircraft being controlled as a function of the calculated difference.
17. A data carrier comprising storing means for storing a computer program product comprising program commands that when executed by a processor of a computer control the computer to perform steps comprising:
detecting aeroelastic and flight-mechanical momenta of an aircraft, positions and movements of control surfaces of the aircraft, and speeds of gusts of wind acting on the aircraft, to generate sensor data;
calculating characteristic quantities of passenger comfort and momenta of the aircraft as a function of the generated sensor data and a stored non-linear simulation model of the aircraft, and
calculating a difference between precalculated quantities and the quantities calculated by the calculation unit, the aircraft being controlled as a function of the calculated difference.
Description This application is a continuation of PCT/EP2009/056659 filed May 29, 2009 and claims the benefit of German Patent Application No. 10 2008 002 124.5 filed May 30, 2008, and U.S. Provisional Application Ser. No. 61/130,375 filed May 30, 2008, the entire disclosures of which are herein incorporated by reference. The invention relates to a calculation system for an aircraft and to a method for determining characteristic quantities and momenta of an aircraft. Aircraft, such as aeroplanes or helicopters, are exposed to various forces during flight. Significant influencing variables in this respect are the lifting forces generated by the aerofoils, the aerodynamic resistance of the aircraft, the weight or gravitational force acting on a centre of gravity of the aircraft, the shear force generated by the engines, the stick forces generated on the control surfaces of the aircraft and the torques caused by the respective forces. The mass inertia of the aircraft or the mass inertia of the aircraft components also plays a part in the above-mentioned forces. Flight maneuvers and air turbulence result in structural loads on the aircraft. To predict the flight behaviour of an aircraft, systems of equations are used which are complex due to the large number of correlations between aeroelastic and flight-mechanical momenta. Conventional simulation systems for simulating the behaviour of aircraft are based on substantially linear models of the structural dynamics of the stationary and instationary aerodynamics, of aeroelastics, of the loads and the flight mechanics. In this respect, conventional systems of equations consider substantially linear characteristics of parameters. The calculation accuracy of these conventional calculation systems using substantially linear models is thus, however, relatively poor, i.e. they do not reflect the actual behaviour of an aircraft in a sufficiently accurate manner. It is therefore an object of the present invention to provide a calculation system and a method for determining characteristic quantities of an aircraft which accurately simulates the actual behaviour of an aircraft. This object is achieved by a calculation system proposed by the present invention. The invention provides a calculation system for an aircraft with at least one sensor for detecting aeroelastic and flight-mechanical momenta of the aircraft, for detecting positions and movements of control surfaces of the aircraft or for detecting speeds of gusts of wind acting on the aircraft; and with a calculation unit which calculates characteristic quantities of passenger comfort and cabin safety as well as momenta of the aircraft as a function of the sensor data provided by the sensors and a non-linear simulation model of the aircraft. In an embodiment of the calculation system according to the invention, the calculation unit automatically adapts the non-linear simulation model using the sensor data provided by the sensors. In an embodiment of the calculation system according to the invention, sensors are provided for detecting momenta of an on-board system of the aircraft. In an embodiment of the calculation system according to the invention, the on-board system has at least one movable mass for damping an associated part of the aircraft. In an embodiment of the calculation system according to the invention, sensors for detecting flight-mechanical momenta of the aircraft also measure deformations of parts of the aircraft. In an embodiment of the calculation system according to the invention, the sensors for detecting flight-mechanical momenta of the aircraft and for detecting aeroelastic momenta of the aircraft have acceleration or pressure sensors. In an embodiment of the calculation system according to the invention, the calculation unit is provided in the aircraft or the data from the sensors of the aircraft is received by the calculation unit via a wireless air interface. In an embodiment of the calculation system according to the invention, the linear simulation model of the aircraft can be read out of a memory. In an embodiment of the calculation system according to the invention, the calculation unit is connected to an input unit for inputting parameters of the simulation model of the aircraft. In an embodiment of the calculation system according to the invention, the calculation unit is connected to an output unit for outputting characteristic quantities and momenta. In an embodiment of the calculation system according to the invention, the on-board system of the aircraft is automatically controlled as a function of the characteristic quantities and momenta, calculated by the calculation unit, for minimising load forces and vibrations. In an embodiment of the calculation system according to the invention, the on-board system of the aircraft can be connected to and disconnected from different frequency ranges. In an embodiment of the calculation system according to the invention, various masses of the on-board system which are fitted to parts of the aircraft can be activated as a function of an adjustable operating mode of the on-board system. The invention also provides a method for determining characteristic quantities of passenger comfort and momenta of an aircraft, comprising the following steps: -
- (a) detection of aeroelastic and flight-mechanical momenta of the aircraft, of positions and movements of control surfaces of the aircraft and of speeds of gusts of wind acting on the aircraft, to generate sensor data; and
- (b) calculation of the characteristic quantities of passenger comfort and of momenta of the aircraft as a function of the generated sensor data and a stored, non-linear simulation model of the aircraft.
The invention also provides a computer program with program commands for implementing a method for determining characteristic quantities of passenger comfort and momenta of an aircraft, which comprises the following steps: -
- (a) detection of aeroelastic and flight-mechanical momenta of the aircraft, of positions and movements of control surfaces of the aircraft and of speeds of gusts of wind acting on the aircraft, to generate sensor data; and
- (b) calculation of the characteristic quantities of passenger comfort and of momenta of the aircraft as a function of the generated sensor data and a stored, non-linear simulation model of the aircraft.
The invention also provides a data carrier which stores such a computer program. In the following, preferred embodiments of the calculation system according to the invention and of the method according to the invention for determining characteristic quantities of passenger comfort and of momenta of an aircraft will be described with reference to the accompanying figures to illustrate features which are essential to the invention. As can be seen from The movement of a rigid aircraft can be described by parameters. In each case three of these variables are combined into a vector, describing the
The causes of the movement are the forces acting on the aircraft,
A further important quantity is the specific force measured by the accelerometers.
The specific force is an indication of the acceleration impression of the pilot according to magnitude and direction and is defined as the ratio of the resulting external force to the mass of the aircraft. To determine Newton's equations and the angular momentum equation, the accelerations and speeds are measured with respect to an inertial system. The earth is used as the inertial system, an earth frame of reference F To simply describe the aerodynamic forces, an aerodynamic coordinate system F The transition from the body frame to the earth frame of reference is made using a transformation matrix L
The subscript index indicates the coordinate system in which the vectors are presented. For example, the vector {right arrow over (R)} _{EB} {right arrow over (R)} _{B} (10)To simplify the notation, the index B is omitted in the following if it is not absolutely necessary. When considering speed, a distinction must also be made between wind and calm. The following generally applies with the speed addition law:
With the components of the vectors {right arrow over (V)}, {right arrow over (Ω)}, and {right arrow over (Φ)} as state quantities, the motion equations are obtained in state space during a calm from Newton's equation and the angular momentum equation, as well as from the relationship between the Euler angles and the rates thereof. The equations apply in particular when the earth is considered as an inertial system with a homogeneous gravitational field and the aeroplane or aircraft is symmetrical with respect to its x-z plane. The arising forces engage according to the model in the centre of gravity and the generation of aerodynamic forces is quasi-stationary. Newton's equation for the centre of gravity of the aircraft in earth frame coordinates is:
Using the transformation matrix
The following applies:
_{EB} {right arrow over (V)}= L _{EB}({right arrow over (Ω)}×{right arrow over (V)}) (14)from which follows: L _{EB} {right arrow over (F)}= L _{EB } m({right arrow over (Ω)}×{right arrow over (V)}+{right arrow over ({dot over (V)})}) (15)The resulting force {right arrow over (F)} is composed of the aerodynamic force {right arrow over (R)} and the weight force {right arrow over (G)}=
Thus, the equations for the speeds are established. The relationships for the rates are obtained analogously from the angular momentum equation with the angular momentum {right arrow over (H)} and the inertia sensor I.
These relationships, divided up into components, produce together with the equations between Euler angles and the rates thereof, the state equations of a rigid aircraft.
By transforming the speed {right arrow over (V)} into the earth frame of reference, where
_{EB} {right arrow over (V)} (19)the differential equations are obtained for calculating the position:
For the specific force, the following is obtained in body frame coordinates for a sensor located on the x-axis at a spacing x
If the vector entries are divided by the gravitational acceleration The above motion equations apply ideally to a rigid aircraft. However, in practice elastic deformations of the structure occur which have a significant influence on the dynamic characteristics of the system. Therefore, the model is expanded by these elastic degrees of freedom. Quasi-static deformations are provided when the natural frequencies of the elastic modes are substantially higher than those of the rigid body modes. In this case, the influence of elastic deformation can be considered by a corresponding adaption of the aerodynamic derivatives. If the natural frequencies of the elastic degrees of freedom are within the same range, the movement of the rigid body is influenced by the elastic deformations. In this case, the dynamics of the elastic degrees of freedom are to be considered in the motion equations. For this purpose, the deformation of the structure can be approximately described by the superposition of normal modes of the free vibration:
x′, y′ z′ are the deflections of the respective rest positions x
The approximation consists in disregarding all couplings over the damping terms between the individual modes. On the assumption that the influence of the degrees of freedom of the rigid body on the elastic modes can be described by a linear correlation and that the elastic deformations are adequately small, the generalised force F
The infinite series occurring here can be replaced by finite series which only retain those modes which lie in the range of the rigid body frequencies. It can be assumed for the further calculation that these are K modes which are combined in a vector
To arrive at a compact notation for all modes, the generalised moments of inertia I ^{T}{dot over (ε)}+ω ω ^{T} ε=I^{−1}( a _{u} Δu+ a _{{dot over (u)}} {dot over (u)}+ . . . + a _{p} p+ . . . + a _{δ} _{ r }δ_{r} + . . . + A _{ε} ε+B _{{dot over (ε)}} {dot over (ε)}+C _{ ε }{umlaut over (ε)}) (26)Presentation in the state space is achieved by introducing the mode speed {dot over ( ^{T} υ+ω ω ^{T} ε=I^{−1}( a _{u} Δu+ a _{{dot over (u)}} {dot over (u)}+ . . . + a _{p} p+ . . . + a _{δ} _{ r }δ_{r}+ . . . . + A _{ε} ε+ B _{υ} υ+ C _{{dot over (υ)}}{dot over (υ)}) (27)Using the Matrices
_{{dot over (x)}} _{ 1 }=[a _{{dot over (u)}} a _{{dot over (υ)}} a _{{dot over (w)}} a _{{dot over (p)}} a _{{dot over (q)}} a _{{dot over (r)}}], A _{x} _{ 1 }=[a _{u } a _{υ} a _{w } a _{p } a _{q } a _{r}], A _{c}=[a _{δ} _{ E } a _{δ} _{ A } a _{δ} _{ R } a _{δ} _{ c } a _{δ} _{ F }], (28)and the unit matrix of k order I, the state equation can be formulated:
_{k} {dot over (ε)}= υ {dot over ( υ)}=( I _{k} − I ^{−1} C _{{dot over (υ)}})^{−1}[( I ^{−1} B _{υ}−2 d ω ^{T})υ+( I ^{−1} A _{ε}−ω ω ^{T})ε+ A _{{dot over (x)}} _{ 1 } x _{1} + A _{x} _{ 1 } x _{1} + A _{c} (29)c]The external forces acting on an aircraft are, in addition to the weight, the aerodynamic forces of lift and resistance as well as thrust. The point of application of the lift is in what is known as the neutral point which is different from the centre of gravity. As a result, moments are generated. This applies similarly to thrust. The resulting forces are combined in a vector {right arrow over (R)} and the moments are combined in a vector {right arrow over (Q)}. Lift and resistance are generated by the relative movement of aircraft and air, i.e. by {right arrow over (V)} and {right arrow over (Ω)}. Furthermore, these forces depend on the angle of attack α and the angles of the control surfaces of the primary flight control, elevator (δ The horizontal symmetrical straight flight can be selected as the stationary flight state. If the stability axes are additionally selected as a flight frame of reference, the above relationships are simplified in that in this state, X
The quantities which occur in equation (31) and are indexed with u and ε describe the influence of the elastic modes on the aerodynamics. They are in each case vectors of length k, with k being the number of elastic modes. The derivatives indexed with c are also vectors which describe the influence of the control factors. The dimension thereof is equal to the number of control factors. The equations derived above are combined into a model by which the entire dynamics of the flexible aircraft can be described under the conditions mentioned in the preceding paragraphs. The states for describing the movement of the rigid body are combined in the vector
_{1} =[Δu w q θ υ p r φ ψ] ^{T} (32)
The partial matrices used in equations (33) and (34) are compiled with the following abbreviations:
The matrices B are obtained by respectively replacing the index ε in the matrix A_{1} _{12 }by u and c. The following applies to the remaining matrices:
The matrices D are obtained by replacing the index ε by u and c, respectively. H and h(x are as follows:_{1})
The non-linear simulation model described in equation (33) contains an effectiveness matrix F which considers the non-linear characteristics of parameters. The effectiveness matrix F is stated in equation (42). Expanding the model by aerodynamic, structural dynamic and aeroelastic non-linearities produces a.) additional entries in the non-linearity vector g(x1),for example g b.) additional columns in the
The quantities X The non-linear simulation model presented in equation (33) can also be described in a physically more concrete manner (in a generalisation of the Newton and Euler motion equations) as follows:
Since the transformation of equation (33) into the form of equation (50) results in modified vectors x and g (x, {dot over (x)}, p, t) and a modified matrix F, these new vectors and matrices are not underlined. The equation system is illustrated graphically in the diagram of Positioned on the right-hand side of the equation system is a hyper-input vector p of the aircraft with a plurality of subvectors. The vector x forms a hyper-momentum vector of the aircraft. The second time derivative x Further non-linear expansions can be presented very graphically in this illustration. Additional non-linearities in the engine dynamics, in the system behavior or in the case of faults expand the non-linearity vector g (x, {dot over (x)}, p, t) and the effectiveness matrix F in additional entries. The matrix entries of F in turn describe the influence strength of non-linearities, as an effective force or moment in the generalised Newton and Euler motion equations. The mass matrix M, the damping matrix D and the rigidity matrix K are expanded matrices which consider the aerodynamics. In the special case shown in In the special case shown in In the special case shown in For the special case shown in In a possible embodiment, the calculation unit The calculation system In the embodiment shown in The calculation unit Furthermore, the calculation system This can also reduce operating costs for the customer and operator of the aircraft In the calculation system In a possible embodiment of the method according to the invention, proceeding from a starting model which, for example corresponds to curve I in Patent Citations
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